1. Field of the Invention
This invention relates to gas turbine engines and more particularly to engines having a shroud surrounding the tips of the rotor blades in the turbine section of the engine.
2. Description of the Prior Art
In a gas turbine engine or the type referred to above, pressurized air and fuel are burned in a combustion chamber to add thermal energy to the medium gases flowing therethrough. The effluent from the chamber comprises high temperature gases which are flowed downstream in an annular flow path through the turbine section of the engine. Nozzle guide vanes at the inlet to the turbine direct the medium gases onto a multiplicity of blades which extend radially outward from the engine rotor. An annular shroud which is supported by the turbine case surrounds the tips of the rotor blades to confine the medium gases flowing thereacross to the flow path. The clearance between the blade tips and the shroud is minimized to prevent the leakage of medium gases around the tips of the blades.
A limiting factor in many turbine engine designs is the maximum temperature of the medium gases which can be tolerated in the turbine without adversely limiting the durability of the individual components. The shrouds which surround the tips of the rotor blades are particularly susceptible to thermal damage and a variety of cooling techniques is applied to control the temperature of the material comprising the shroud in the face of high turbine inlet temperatures. In many of these techniques air is bled from the compressor through suitable conduit means to the local area to be cooled. Compressor air is sufficiently high in pressure to cause the air to flow into the local area of the turbine without auxiliary pumping and is sufficiently low in temperature to provide the required cooling capacity.
Most recently, considerable design effort has been expended to minimize the amount of air consumed for cooling of the turbine components. Impingement cooling is one of the more effective techniques utilized and occurs where a high velocity air stream is directed against a component to be cooled. The high velocity stream impinges upon a surface of the component and increases the rate of heat transfer between the component and the cooling air. A second highly effective but not as widely utilized technique is that of transpiration cooling. A cooling medium is allowed to exude at low velocities through a multiplicity of tiny orifices in the wall of the component to be cooled. The low velocity flow adheres to the external surface of the component and is carried axially downstream along the surface by the working medium gases flowing thereacross.
One typical application of transpiration cooling to blade tip shrouds is shown in U.S. Pat. No. 3,365,175 to McDonough et al. entitled "Air Cooled Shroud Seal" . In McDonough et al. a single cooling air chamber extends circumferentially about the outer periphery of the shroud. Cooling air is flowable to the chamber from the compressor section of the engine through suitable supply means to convectively cool the shroud material. At least a portion of the cooling air in McDonough et al. is further flowable to the inner periphery of the shroud through cooling holes of small diameter to introduce cool air into the boundary layer of the hot gas stream adjacent the shroud. One embodiment of McDonough et al, has a multiplicity of grooves or recesses at the inner periphery of the shroud which intercept the cooling holes and prevent the closure of the holes should the blade tips rub against the shroud during operation of the engine. In transpiration cooling the exuding velocities must remain low in order to prevent over penetration of the working medium gases by the cooling air. Over penetration interrupts both the flow of cooling air and the flow of medium gases and renders the cooling ineffective. A preferred pressure ratio across the cooled wall in most transpiration cooled embodiments is approximately 1.25. The effectiveness of a transpiration cooled construction is highly sensitive to variations from the designed pressure ratio across the surface to be cooled; accordingly, the pressure ratio must be closely controlled.
Both cooled and uncooled shrouds are commonly segmented where large variations in thermal expansion between the shroud and its supporting turbine case are expected. A circumferential gap between adjacent segments is provided to allow independent expansion of the case and shroud segments without inducing local stresses. In this type of construction a portion of the medium gases inherently leaks axially through the gap from the upstream to the downstream region of the shroud. A reduction in the amount of leaking gases is effected by providing interlocking lugs at the abutting ends of adjacent segments. U.S. Pat. No. 3,412,977 to Moyer et al. entitled "Segmented Annular Sealing Ring and Method of its Manufacture" shows a shroud having conventionally interlocking lugs. In addition to the interlocking lugs, shroud constructions which are both segmented and cooled require radial sealing means to prevent the wasteful leakage of cooling air from the air chamber into the medium flow path through the gap between adjacent segments. To be effective the radial sealing means must necessarily have a capability for sealing a gap which varies in width according to divergent thermal conditions.
The individual use of the above described cooling techniques and sealing means, although successful in prolonging the life of the turbine components, have proved inadequate to meet today's requirement for durable, high performance engines. More effective ways of utilizing a diminished quantity of cooling air must be found.